Gas turbine engines are generally known for use in a wide range of applications such as aircraft engines and auxiliary power units for aircraft. In a typical configuration, the gas turbine engine includes a turbine section having a plurality of sets or rows of stator vanes and turbine blades disposed in an alternating sequence along an axial length of a hot gas flow path of generally annular shape. The turbine blades are coupled to a main engine shaft through one or more rotor disks. Hot combustion gases are delivered from an engine combustor to the annular hot gas flow path, resulting in rotary driving of the turbine rotor disks which, in turn, drives the compressors and gearbox.
Advanced high performance gas turbine engines are constantly driven to achieve maximized thermodynamic efficiency, which is generally achieved by operating at higher rotor speeds and temperatures. In many gas turbine engine configurations the turbine blades are mounted at the periphery of the one or more rotor disks through a mechanical connection, e.g., through a dovetail-type connection or the like. However, the mechanical properties of the rotor disks and turbine blades may be inadequate to sustain induced loads during operation, even with selection of special materials and engineered cooling schemes. This may be especially true as efforts are made to maximize thermodynamic efficiency by maximizing rotor speeds and operating temperatures.
One approach taken to maximize temperatures and load carrying capability in turbine blades and rotor disks, particularly in the high pressure turbine (HPT) section, is to employ dissimilar materials for the rotor disks and the turbine blades while removing the stress concentrations associated to mechanical connections. The respective rotor disks and turbine blades, including the dissimilar materials, are directly bonded together as opposed to relying upon a mechanical connection. In one example, the turbine blades may be operatively connected to blade mounts, e.g., by casting the turbine blades and blade mounts together, or by brazing or welding the turbine blades to the blade mounts. The blade mounts may be operatively connected to each other forming a blade ring, such as by casting a plurality of blade mounts together or by brazing or welding blade mounts together. The creation of an integral bonded rotor requires the release of hoop stress attributable to the thermal gradients and rotation of the rotor disk. The hoop stress can be broken by slotting the blade ring and rotor disk after bonding the blade ring and rotor disk together.
With bonded turbine blade/rotor disk configurations, radial stress at a bond line between the blade ring and the rotor disk is often a concern and can lead to structural failure of the bonded turbine blade/rotor disk. Various bond line geometries have been proposed to address assembly and bonding surface area considerations. However, radial stress is still a concern with such bond line geometries. Further, various configurations of turbine wheels exhibit areas of higher radial stress along the bond line due to particular features of the turbine wheels.
Accordingly, it is desirable to provide turbine wheels, turbine engines including the turbine wheels, and methods of fabricating the turbine wheels having improved bond line geometry for minimizing radial stress at the bond line. Furthermore, other desirable features and characteristics will become apparent from the subsequent detailed description and the appended claims, taken in conjunction with the accompanying drawings and this background.